EP3572622B1 - Carter intermédiaire de turbine pourvu de contour d'espace annulaire à conception spécifique - Google Patents

Carter intermédiaire de turbine pourvu de contour d'espace annulaire à conception spécifique Download PDF

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Publication number
EP3572622B1
EP3572622B1 EP19175531.3A EP19175531A EP3572622B1 EP 3572622 B1 EP3572622 B1 EP 3572622B1 EP 19175531 A EP19175531 A EP 19175531A EP 3572622 B1 EP3572622 B1 EP 3572622B1
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EP
European Patent Office
Prior art keywords
annular space
wall
trailing edge
leading edge
turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP19175531.3A
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German (de)
English (en)
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EP3572622A1 (fr
Inventor
Markus Brettschneider
Christoph Lauer
Rudolf Stanka
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MTU Aero Engines AG
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MTU Aero Engines AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
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Publication date
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Publication of EP3572622A1 publication Critical patent/EP3572622A1/fr
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/713—Shape curved inflexed
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a turbine center frame for a gas turbine, in particular an aircraft gas turbine, with a radial inner wall, a radial outer wall, the inner wall and the outer wall delimiting an annular space through which hot gas flows, and the inner wall and the outer wall each having a contour facing the annular space, which describe an inner annulus curve along the inner wall and an outer annulus curve along the outer wall with respect to an axial longitudinal section through the turbine center frame, and having at least one blade element which extends through the annulus in the radial direction and has an axial leading edge and an axial trailing edge, the The vane element has an outer axial width relative to the outer wall measured between the leading edge and the trailing edge and an inner axial width relative to the inner wall measured between the leading edge and the trailing edge.
  • the vane element of such a turbine center frame may be a turning or non-turning vane.
  • a deflecting blade is understood to be a blade element that has a significant influence on the direction of flow of hot gas flowing through the annular space and is not just flowed around without the direction of flow being significantly influenced.
  • the blade elements distributed along the circumferential direction serve to surround supporting structures that extend from a hub area to a housing area of a gas turbine through the annular space in a streamlined manner and to protect them from the hot gas flowing through.
  • An example of the design of an annular space of a turbine center frame is from US 2014/0086739 A1 known. More generic turbine center frame are also from DE 10 2004 042699 A1 and the U.S. 2010/040462 A1 known.
  • the outer wall and inner wall have a steeper gradient. There is a risk that flow separation can occur in the area of the outer wall or the inner wall.
  • the object on which the invention is based is seen in specifying a turbine center frame in which a minimum axial length can be realized and the risk of flow separations along the walls delimiting the annular space is minimized.
  • the provision of turning points or points of maximum gradient in the areas of the inlet edge or the outlet edge enables an optimized design of the annular space curves with regard to a shortened axial length.
  • the provision of the turning points or the points of maximum gradient in the area of the blade edges also prevents the flow from breaking away along the annular space walls.
  • In the area of the leading edge or the trailing edge there is at least one turning point of the outer annular space curve or the inner annular space curve.
  • the turning point of an annulus curve is understood to be that point along the annulus curve at which the curvature of the annulus curve changes from convex to concave or vice versa, it being possible for the designations convex and concave to be specified in relation to a hub of the turbine center frame.
  • the projected length of the curve portion may include a front portion forward of the leading edge or the trailing edge and a back portion after the leading edge or the trailing edge, the front portion and the back portion being of substantially equal length.
  • the front section and the rear section have the same lengths in relation to the penetration point of the leading edge or trailing edge through the inner wall or the outer wall.
  • the front section and the rear section have a length which is up to 10% of the inner or outer axial width of the blade element.
  • the outer annular space curve has a turning point in the region of the leading edge and a turning point in the region of the trailing edge.
  • the inner annulus curve can have an inflection point in the area of the leading edge and an inflection point in the area of the trailing edge. It is therefore possible for the annular space curves to be designed in such a way that their respective turning points, in particular all turning points, are only provided in the area of the leading edge or the trailing edge, in particular in the area of the corresponding curve section with the projected length of 20% of the axial width of the blade element.
  • the point of maximum gradient of the outer annulus curve can be provided in the area of the leading edge or in the area of the trailing edge.
  • the point of maximum gradient of the inner annular space curve can be provided in the area of the trailing edge.
  • an annulus can be designed in such a way that its points of maximum slope are located at three points, all of which are located in the region of the leading edge and the trailing edge. In this case, the inner annular space curve has no point of maximum gradient, particularly at the leading edge.
  • the invention also relates to a gas turbine, in particular an aircraft gas turbine, with at least two consecutive turbines, in particular a high-pressure turbine and a low-pressure turbine or in particular with a high-pressure turbine, an intermediate-pressure turbine and a low-pressure turbine, with between two consecutive turbines, in particular between the high-pressure turbine and the subsequent low-pressure turbine or intermediate-pressure turbine , one described above is installed such that hot gas flowing out of one turbine is passed through the annular space to the following turbine.
  • a gas turbine in particular an aircraft gas turbine, with at least two consecutive turbines, in particular a high-pressure turbine and a low-pressure turbine or in particular with a high-pressure turbine, an intermediate-pressure turbine and a low-pressure turbine, with between two consecutive turbines, in particular between the high-pressure turbine and the subsequent low-pressure turbine or intermediate-pressure turbine, one described above is installed such that hot gas flowing out of one turbine is passed through the annular space to the following turbine.
  • FIG. 1 shows a schematic and simplified view of an aircraft gas turbine 10, which is illustrated purely by way of example as a turbofan engine.
  • the gas turbine 10 includes a fan 12 surrounded by a jacket 14 that is indicated.
  • the fan 12 is followed by a compressor 16, which is accommodated in an indicated inner housing 18 and can be of single-stage or multi-stage design.
  • the compressor 16 is followed by the combustion chamber 20 .
  • Hot exhaust gas flowing out of the combustion chamber flows men and can be designed in one or more stages.
  • the compressor 16 is followed by the combustion chamber 20 .
  • Hot exhaust gas flowing out of the combustion chamber then flows through the adjoining turbine 22, which can be of single-stage or multi-stage design.
  • the turbine 22 includes a high-pressure turbine 24 and a low-pressure turbine 26.
  • a hollow shaft 28 connects the high-pressure turbine 24 to the compressor 16, in particular a high-pressure compressor 29, so that they are driven or rotated together.
  • a further inner shaft 30 in the radial direction RR of the turbine connects the low-pressure turbine 26 to the fan 12 and to a low-pressure compressor 32 here, so that they are driven or rotated together.
  • An outlet housing 33 which is only indicated here, is connected to the turbine 22 .
  • a turbine center frame 34 is arranged between the high-pressure turbine 24 and the low-pressure turbine 26, which is arranged around the shafts 28, 30.
  • hot exhaust gases from the high-pressure turbine 24 flow through the turbine center frame 34 .
  • the hot exhaust gas then reaches an annular space 38 of the low-pressure turbine 26.
  • Rotor blade rings 27 of the compressors 28, 32 and the turbines 24, 26 are shown by way of example.
  • guide vane rings 31 that are usually present are shown as an example only for compressor 32 .
  • figure 2 shows a longitudinal section through the annular space 38 of a turbine center frame 34.
  • the sectional plane is spanned by the axial direction AR and the radial direction RR.
  • the turbine center frame 34 includes a radial inner wall 40 and a radial outer wall 42.
  • the inner wall 40 and the outer wall 42 delimit the annular space 38 through which hot gas flows.
  • the inner wall 40 and the outer wall 42 each have a contour 40a, 42a facing the annular space 38.
  • the two contours 40a, 42a describe an inner annulus curve 44 along the inner wall 40 and an outer annulus curve 46 along the outer wall 42 in relation to the axial longitudinal section through the turbine center frame 34.
  • a vane element 48 can also be seen, which extends through the annular space 38 in the radial direction RR.
  • the vane element 48 has an axial interior axial width AB and an inner axial width IB of the blade element 48 measured between the leading edge 50 and the trailing edge 52 and related to the inner wall 40 . It should be noted that a plurality of vane elements 48 are provided on the turbine center frame along the circumferential direction.
  • the outer annular space curve 46 and/or the inner annular space curve 44 has at least one curved section 44c, 44d, 46c, 46d, which has an inflection point 44w, 46w of the relevant annular space curve 44, 46.
  • the curve section can have a point of maximum gradient 44s, 46s of the annular space curve 44, 46 in question.
  • the curve section 44c, 44d, 46c, 46d is arranged in the region of the leading edge 50 or the trailing edge 52 in relation to the outer axial width AB and/or to the inner axial width IB.
  • the curve section 44c, 44d, 46c, 46f has a length KL projected parallel to the axial direction AR, which is up to 20% of the relevant axial width AB or IB.
  • the curve section 44c, 44d, 46c, 46d in question intersects a penetration point 60 of the leading edge 50 or the trailing edge 52 through the radial outer wall 42 or the radial inner wall 40.
  • the projected length KL of the curve section 44c, 44d, 46c, 46d in question can have a front section KLv lying in front of the leading edge 50 or the trailing edge 52 and a rear section KLh lying after the leading edge 50 or the trailing edge 52, the front section KLv and the rear section KLh are essentially the same length.
  • turning points 44w, 46w and/or points of maximum gradient 44s, 46s are within an area that is a maximum of 10% of the relevant axial width AB or IB at the relevant position (on the housing or hub side) from the respective penetration point 60 .
  • turning points 44w, 46w can be provided simultaneously both in the area of the leading edge 50 and in the area of the trailing edge 52. However, it is pointed out that it is also conceivable that in less than As from the representation of figure 3 As can be seen, turning points 44w, 46w can be provided simultaneously both in the area of the leading edge 50 and in the area of the trailing edge 52. However, it is pointed out that it is also conceivable that a turning point 44w, 46w can be provided in less than the four curve sections 44c, 44d, 46c, 46d shown. In particular, such a turning point can also be arranged in just one of the curve sections 44c, 44d, 46c, 46d. Incidentally, this also applies to the points of maximum gradient 44s, 46s. However, it is pointed out that there is usually no point of maximum gradient 44s of the inner annular space curve 44 in the curve section 44c.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (5)

  1. Carter intermédiaire de turbine pour une turbine à gaz, en particulier une turbine à gaz d'aéronef, comportant
    une paroi interne (40) radiale ;
    une paroi externe (42) radiale ;
    dans lequel la paroi interne (40) et la paroi externe (42) délimitent un espace annulaire (38) traversé par du gaz chaud et dans lequel la paroi interne (40) et la paroi externe (42) comportent un contour (40a, 42a) respectif faisant face à l'espace annulaire (38), lequel décrit, par rapport à une coupe longitudinale axiale à travers le carter intermédiaire de turbine (34), une courbe d'espace annulaire (44) intérieure le long de la paroi interne (40) et une courbe d'espace annulaire (46) extérieure le long de la paroi externe (42) ;
    au moins un élément d'aube (48), lequel s'étend à travers l'espace annulaire (38) dans la direction radiale (RR) et comporte un bord d'attaque (50) axial et un bord de fuite (52) axial, dans lequel l'élément d'aube (48) présente une largeur axiale (AB) extérieure mesurée entre le bord d'attaque (50) et le bord de fuite (52) par rapport à la paroi externe (42) et une largeur axiale (IB) intérieure mesurée entre le bord d'attaque (50) et le bord de fuite (52) par rapport à la paroi interne (40),
    dans lequel la courbe d'espace annulaire (46) extérieure et/ou la courbe d'espace annulaire (44) intérieure comportent au moins une section de courbe (44c, 44d, 46c, 46d), laquelle présente un point de pente maximale (44s, 46s) de la courbe d'espace annulaire (44, 46) concernée, dans lequel la section de courbe (44c, 44d, 46c, 46d) se trouve, par rapport à la largeur axiale (AB) extérieure et/ou la largeur axiale (IB) intérieure, dans la zone du bord d'attaque (50) ou du bord de fuite (52) et dans lequel la section de courbe (44c, 44d, 46c, 46d) coupe le point de percée (60) du bord d'attaque (50) ou du bord de fuite (52) à travers la paroi externe (42) ou la paroi interne (40), caractérisé en ce que la section de courbe (44c, 44d, 46c, 46d) présente une longueur projetée (KL) parallèle à la direction axiale (AR), laquelle atteigne jusqu'à 20 % de la largeur axiale (AB, IB) concernée et la courbe d'espace annulaire (46) extérieure présente un point d'inflexion (46w) dans la zone du bord d'attaque (50) et un point d'inflexion (46w) dans la zone du bord de fuite (52).
  2. Carter intermédiaire de turbine selon la revendication 1, caractérisé en ce que la longueur projetée (KL) de la section de courbe (44c, 44d, 46c, 46d) comporte une section avant (KLv) située devant le bord d'attaque (50) ou le bord de fuite (52) et une section arrière (KLh) située derrière le bord d'attaque (50) ou le bord de fuite (52), dans lequel la section avant (KLv) et la section arrière (KLh) sont sensiblement de la même longueur, et/ou que la courbe d'espace annulaire (46) extérieure et/ou la courbe d'espace annulaire (44) intérieure comportent au moins une section de courbe (44c, 44d, 46c, 46d), laquelle présente un point d'inflexion (44w, 46w) de la courbe d'espace annulaire (44, 46) concernée, dans lequel la section de courbe (44c, 44d, 46c, 46d) se trouve dans la zone du bord d'attaque (50) ou du bord de fuite (52) par rapport à la largeur axiale (AB) extérieure et/ou la largeur axiale (IB) intérieure et présente une longueur projetée (KL) parallèle à la direction axiale (AR) et atteignant jusqu'à 20 % de la largeur axiale (AB, IB) concernée et dans lequel la section de courbe (44c, 44d, 46c, 46d) coupe le point de percée (60) du bord d'attaque (50) ou du bord de fuite (52) à travers la paroi externe (42) ou la paroi interne (40).
  3. Carter intermédiaire de turbine selon l'une quelconque des revendications précédentes, caractérisé en ce que la courbe d'espace annulaire (44) intérieure présente un point d'inflexion (44w) dans la zone du bord d'attaque (50) et un point d'inflexion (44w) dans la zone du bord de fuite (52).
  4. Carter intermédiaire de turbine selon l'une quelconque des revendications précédentes, caractérisé en ce que le point de la pente maximale (46s) de la courbe d'espace annulaire (46) extérieure est situé dans la zone du bord d'attaque (50) ou dans la zone du bord de fuite (52), et/ou que le point de la pente maximale (44s) de la courbe d'espace annulaire (44) intérieure est situé dans la zone du bord de fuite (52).
  5. Turbine à gaz, en particulier turbine à gaz d'aéronef, comportant au moins deux turbines (24, 26) se succédant, en particulier une turbine haute pression et une turbine basse pression ou, en particulier comportant une turbine haute pression, une turbine moyenne pression et une turbine basse pression, dans laquelle un carter intermédiaire de turbine (34) selon l'une quelconque des revendications précédentes est installé entre deux turbines (24, 26) se succédant de telle sorte que du gaz chaud sortant d'une turbine (24) est guidé à travers l'espace annulaire (38) jusqu'à la turbine (26) successive.
EP19175531.3A 2018-05-24 2019-05-21 Carter intermédiaire de turbine pourvu de contour d'espace annulaire à conception spécifique Active EP3572622B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102018208151.4A DE102018208151A1 (de) 2018-05-24 2018-05-24 Turbinenzwischengehäuse mit spezifisch ausgebildeter Ringraumkontur

Publications (2)

Publication Number Publication Date
EP3572622A1 EP3572622A1 (fr) 2019-11-27
EP3572622B1 true EP3572622B1 (fr) 2022-11-23

Family

ID=66625811

Family Applications (1)

Application Number Title Priority Date Filing Date
EP19175531.3A Active EP3572622B1 (fr) 2018-05-24 2019-05-21 Carter intermédiaire de turbine pourvu de contour d'espace annulaire à conception spécifique

Country Status (4)

Country Link
US (1) US10876418B2 (fr)
EP (1) EP3572622B1 (fr)
DE (1) DE102018208151A1 (fr)
ES (1) ES2933927T3 (fr)

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GB2267736B (en) * 1992-06-09 1995-08-09 Gen Electric Segmented turbine flowpath assembly
US5397215A (en) 1993-06-14 1995-03-14 United Technologies Corporation Flow directing assembly for the compression section of a rotary machine
DE19650656C1 (de) 1996-12-06 1998-06-10 Mtu Muenchen Gmbh Turbomaschine mit transsonischer Verdichterstufe
DE102004042699A1 (de) * 2004-09-03 2006-03-09 Mtu Aero Engines Gmbh Strömungsstruktur für eine Gasturbine
US7594388B2 (en) * 2005-06-06 2009-09-29 General Electric Company Counterrotating turbofan engine
US7870719B2 (en) * 2006-10-13 2011-01-18 General Electric Company Plasma enhanced rapidly expanded gas turbine engine transition duct
US8061980B2 (en) 2008-08-18 2011-11-22 United Technologies Corporation Separation-resistant inlet duct for mid-turbine frames
DE102008060847B4 (de) * 2008-12-06 2020-03-19 MTU Aero Engines AG Strömungsmaschine
US20120275922A1 (en) * 2011-04-26 2012-11-01 Praisner Thomas J High area ratio turbine vane
US9534497B2 (en) * 2012-05-02 2017-01-03 Honeywell International Inc. Inter-turbine ducts with variable area ratios
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EP3064706A1 (fr) 2015-03-04 2016-09-07 Siemens Aktiengesellschaft Rangée d'aubes directrices pour une turbomachine traversée axialement
EP3159505B1 (fr) * 2015-10-20 2020-01-08 MTU Aero Engines GmbH Carter intermédiaire pour une turbine a gaz
US10344602B2 (en) * 2016-04-18 2019-07-09 General Electric Company Gas turbine engine transition duct and turbine center frame
US20180306041A1 (en) * 2017-04-25 2018-10-25 General Electric Company Multiple turbine vane frame

Also Published As

Publication number Publication date
DE102018208151A1 (de) 2019-11-28
ES2933927T3 (es) 2023-02-14
EP3572622A1 (fr) 2019-11-27
US20190360347A1 (en) 2019-11-28
US10876418B2 (en) 2020-12-29

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